Gas turbine engine compressor arrangement

ABSTRACT

A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low pressure compressor with four (4) stages.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a Continuation of U.S. patent application Ser.No. 12/131,876, filed Jun. 2, 2008 U.S. Pat. No. 8,128,021.

BACKGROUND

The present invention relates to a gas turbine engine and moreparticularly to an engine mounting configuration for the mounting of aturbofan gas turbine engine to an aircraft pylon.

A gas turbine engine may be mounted at various points on an aircraftsuch as a pylon integrated with an aircraft structure. An enginemounting configuration ensures the transmission of loads between theengine and the aircraft structure. The loads typically include theweight of the engine, thrust, aerodynamic side loads, and rotary torqueabout the engine axis. The engine mount configuration must also absorbthe deformations to which the engine is subjected during differentflight phases and the dimensional variations due to thermal expansionand retraction.

One conventional engine mounting configuration includes a pylon having aforward mount and an aft mount with relatively long thrust links whichextend forward from the aft mount to the engine intermediate casestructure. Although effective, one disadvantage of this conventionaltype mounting arrangement is the relatively large “punch loads” into theengine cases from the thrust links which react the thrust from theengine and couple the thrust to the pylon. These loads tend to distortthe intermediate case and the low pressure compressor (LPC) cases. Thedistortion may cause the clearances between the static cases androtating blade tips to increase which may negatively affect engineperformance and increase fuel burn.

SUMMARY

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes a fan section, a low spool that includes a lowpressure compressor section, a high spool that includes a high pressurecompressor section, a gear train defined along an engine centerlineaxis, the low spool operable to drive the fan section through said geartrain, an overall pressure ratio provided by the combination of the lowpressure compressor section and the high pressure compressor section,the low pressure compressor section includes four (4) stages, and thehigh pressure compressor section includes eight (8) stages to providethe overall pressure ratio.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low spool may include a three-stage low pressureturbine.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than about five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan bypass flow may define a bypass ratiogreater than ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gear train may define a gear reduction ratio ofgreater than or equal to 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the engine may include a fan variable area nozzle tovary a fan nozzle exit area and adjust a pressure ratio of a fan bypassairflow of the fan section during engine operation.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the engine may include a controller operable tocontrol the fan variable area nozzle to vary a fan nozzle exit area andadjust the pressure ratio of the fan bypass airflow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the controller may be operable to reduce the fannozzle exit area at a cruise flight condition.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the controller is operable to control the fan nozzleexit area to reduce a fan instability.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan variable area nozzle may define a trailingedge of the fan nacelle.

A gas turbine engine according to another exemplary aspect of thepresent disclosure includes a gear train defined along an enginecenterline axis, the gear train defines a gear reduction ratio ofgreater than or equal to 2.5, and a spool along the engine centerlineaxis which drives the gear train, the spool includes a three-stage lowpressure turbine and a four-stage low pressure compressor.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than five (5). Additionally or alternatively, the bypassflow may define a bypass ratio greater than ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gear train may define a gear reduction ratio ofgreater than or equal to 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than five (5), and the bypass flow may define a bypassratio greater than ten (10).

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently disclosed embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1A is a general schematic sectional view through a gas turbineengine along the engine longitudinal axis;

FIG. 1B is a general sectional view through a gas turbine engine alongthe engine longitudinal axis illustrating an engine static structurecase arrangement on the lower half thereof;

FIG. 1C is a side view of an mount system illustrating a rear mountattached through an engine thrust case to a mid-turbine frame between afirst and second bearing supported thereby;

FIG. 1D is a forward perspective view of an mount system illustrating arear mount attached through an engine thrust case to a mid-turbine framebetween a first and second bearing supported thereby;

FIG. 2A is a top view of an engine mount system;

FIG. 2B is a side view of an engine mount system within a nacellesystem;

FIG. 2C is a forward perspective view of an engine mount system within anacelle system;

FIG. 3 is a side view of an engine mount system within another frontmount;

FIG. 4A is an aft perspective view of an aft mount;

FIG. 4B is an aft view of an aft mount of FIG. 4A;

FIG. 4C is a front view of the aft mount of FIG. 4A;

FIG. 4D is a side view of the aft mount of FIG. 4A;

FIG. 4E is a top view of the aft mount of FIG. 4A;

FIG. 5A is a side view of the aft mount of FIG. 4A in a first slideposition; and

FIG. 5B is a side view of the aft mount of FIG. 4A in a second slideposition.

DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT

FIG. 1A illustrates a general partial fragmentary schematic view of agas turbofan engine 10 suspended from an engine pylon 12 within anengine nacelle assembly N as is typical of an aircraft designed forsubsonic operation.

The turbofan engine 10 includes a core engine within a core nacelle Cthat houses a low spool 14 and high spool 24. The low spool 14 includesa low pressure compressor 16 and low pressure turbine 18. The low spool14 drives a fan section 20 connected to the low spool 14 either directlyor through a gear train 25.

The high spool 24 includes a high pressure compressor 26 and highpressure turbine 28. A combustor 30 is arranged between the highpressure compressor 26 and high pressure turbine 28. The low and highspools 14, 24 rotate about an engine axis of rotation A.

In one disclosed, non-limiting embodiment, the low pressure compressor16 includes 4 stages (16A-16D shown in FIG. 1A), the high pressurecompressor 26 includes 8 stages (26A-26H shown in FIG. 1B) and the lowpressure turbine 18 includes 3 stages (18A-18C shown in FIG. 1B).

The engine 10 in one non-limiting embodiment is a high-bypass gearedarchitecture aircraft engine. In one disclosed embodiment, the engine 10bypass ratio is greater than ten (10:1), the turbofan diameter issignificantly larger than that of the low pressure compressor 16, andthe low pressure turbine 18 has a pressure ratio that is greater than5:1. The gear train 25 may be an epicycle gear train such as a planetarygear system or other gear system with a gear reduction ratio of greaterthan 2.5:1. It should be understood, however, that the above parametersare only exemplary of one embodiment of a geared architecture engine andthat the present invention is applicable to other gas turbine enginesincluding direct drive turbofans.

Airflow enters the fan nacelle F which at least partially surrounds thecore nacelle C. The fan section 20 communicates airflow into the corenacelle C to the low pressure compressor 16. Core airflow compressed bythe low pressure compressor 16 and the high pressure compressor 26 ismixed with the fuel in the combustor 30 where is ignited, and burned.The resultant high pressure combustor products are expanded through thehigh pressure turbine 28 and low pressure turbine 18. The turbines 28,18 are rotationally coupled to the compressors 26, 16 respectively todrive the compressors 26, 16 in response to the expansion of thecombustor product. The low pressure turbine 18 also drives the fansection 20 through gear train 25. A core engine exhaust E exits the corenacelle C through a core nozzle 43 defined between the core nacelle Cand a tail cone 33.

With reference to FIG. 1B, the low pressure turbine 18 includes a lownumber of stages, which, in the illustrated non-limiting embodiment,includes three turbine stages, 18A, 18B, 18C. The gear train 22operationally effectuates the significantly reduced number of stageswithin the low pressure turbine 18. The three turbine stages, 18A, 18B,18C facilitate a lightweight and operationally efficient enginearchitecture. It should be appreciated that a low number of stagescontemplates, for example, 3-5 stages.

The engine static structure 44 generally has sub-structures including acase structure often referred to as the engine backbone. The enginestatic structure 44 generally includes a fan case 46, an intermediatecase (IMC) 48, a high pressure compressor case 50, a combustor case 52A,a high pressure turbine case 52B, a thrust case 52C, a low pressureturbine case 54, and a turbine exhaust case 56 (FIG. 1B). Alternatively,the combustor case 52A, the high pressure turbine case 52B and thethrust case 52C may be combined into a single case. It should beunderstood that this is an exemplary configuration and any number ofcases may be utilized.

The fan section 20 includes a fan rotor 32 with a plurality ofcircumferentially spaced radially outwardly extending fan blades 34. Thefan blades 34 are surrounded by the fan case 46. The core engine casestructure is secured to the fan case 46 at the IMC 48 which includes amultiple of circumferentially spaced radially extending struts 40 whichradially span the core engine case structure and the fan case 20.

The engine static structure 44 further supports a bearing system uponwhich the turbines 28, 18, compressors 26, 16 and fan rotor 32 rotate. A#1 fan dual bearing 60 which rotationally supports the fan rotor 32 isaxially located generally within the fan case 46. The #1 fan dualbearing 60 is preloaded to react fan thrust forward and aft (in case ofsurge). A #2 LPC bearing 62 which rotationally supports the low spool 14is axially located generally within the intermediate case (IMC) 48. The#2 LPC bearing 62 reacts thrust. A #3 fan dual bearing 64 whichrotationally supports the high spool 24 and also reacts thrust. The #3fan bearing 64 is also axially located generally within the IMC 48 justforward of the high pressure compressor case 50. A #4 bearing 66 whichrotationally supports a rear segment of the low spool 14 reacts onlyradial loads. The #4 bearing 66 is axially located generally within thethrust case 52C in an aft section thereof. A #5 bearing 68 rotationallysupports the rear segment of the low spool 14 and reacts only radialloads. The #5 bearing 68 is axially located generally within the thrustcase 52C just aft of the #4 bearing 66. It should be understood thatthis is an exemplary configuration and any number of bearings may beutilized.

The #4 bearing 66 and the #5 bearing 68 are supported within amid-turbine frame (MTF) 70 to straddle radially extending structuralstruts 72 which are preloaded in tension (FIGS. 1C-1D). The MTF 70provides aft structural support within the thrust case 52C for the #4bearing 66 and the #5 bearing 68 which rotatably support the spools 14,24.

A dual rotor engine such as that disclosed in the illustrated embodimenttypically includes a forward frame and a rear frame that support themain rotor bearings. The intermediate case (IMC) 48 also includes theradially extending struts 40 which are generally radially aligned withthe #2 LPC bearing 62 (FIG. 1B). It should be understood that variousengines with various case and frame structures will benefit from thepresent invention.

The turbofan gas turbine engine 10 is mounted to aircraft structure suchas an aircraft wing through a mount system 80 attachable by the pylon12. The mount system 80 includes a forward mount 82 and an aft mount 84(FIG. 2A). The forward mount 82 is secured to the IMC 48 and the aftmount 84 is secured to the MTF 70 at the thrust case 52C. The forwardmount 82 and the aft mount 84 are arranged in a plane containing theaxis A of the turbofan gas turbine 10. This eliminates the thrust linksfrom the intermediate case, which frees up valuable space beneath thecore nacelle and minimizes IMC 48 distortion.

Referring to FIGS. 2A-2C, the mount system 80 reacts the engine thrustat the aft end of the engine 10. The term “reacts” as utilized in thisdisclosure is defined as absorbing a load and dissipating the load toanother location of the gas turbine engine 10.

The forward mount 82 supports vertical loads and side loads. The forwardmount 82 in one non-limiting embodiment includes a shackle arrangementwhich mounts to the IMC 48 at two points 86A, 86B. The forward mount 82is generally a plate-like member which is oriented transverse to theplane which contains engine axis A. Fasteners are oriented through theforward mount 82 to engage the intermediate case (IMC) 48 generallyparallel to the engine axis A. In this illustrated non-limitingembodiment, the forward mount 82 is secured to the IMC 40. In anothernon-limiting embodiment, the forward mount 82 is secured to a portion ofthe core engine, such as the high-pressure compressor case 50 of the gasturbine engine 10 (see FIG. 3). One of ordinary skill in the art havingthe benefit of this disclosure would be able to select an appropriatemounting location for the forward mount 82.

Referring to FIG. 4A, the aft mount 84 generally includes a first A-arm88A, a second A-arm 88B, a rear mount platform 90, a wiffle treeassembly 92 and a drag link 94. The rear mount platform 90 is attacheddirectly to aircraft structure such as the pylon 12. The first A-arm 88Aand the second A-arm 88B mount between the thrust case 52C at casebosses 96 which interact with the MTF 70 (FIGS. 4B-4C), the rear mountplatform 90 and the wiffle tree assembly 92. It should be understoodthat the first A-arm 88A and the second A-arm 88B may alternativelymount to other areas of the engine 10 such as the high pressure turbinecase or other cases. It should also be understood that other framearrangements may alternatively be used with any engine case arrangement.

Referring to FIG. 4D, the first A-arm 88A and the second A-arm 88B arerigid generally triangular arrangements, each having a first link arm 89a, a second link arm 89 b and a third link arm 89 c. The first link arm89 a is between the case boss 96 and the rear mount platform 90. Thesecond link arm 89 b is between the case bosses 96 and the wiffle treeassembly 92. The third link arm 89 c is between the wiffle tree assembly92 rear mount platform 90. The first A-arm 88A and the second A-arm 88Bprimarily support the vertical weight load of the engine 10 and transmitthrust loads from the engine to the rear mount platform 90.

The first A-arm 88A and the second A-arm 88B of the aft mount 84 forcethe resultant thrust vector at the engine casing to be reacted along theengine axis A which minimizes tip clearance losses due to engine loadingat the aft mount 84. This minimizes blade tip clearance requirements andthereby improves engine performance.

The wiffle tree assembly 92 includes a wiffle link 98 which supports acentral ball joint 100, a first sliding ball joint 102A and a secondsliding ball joint 102B (FIG. 4E). It should be understood that variousbushings, vibration isolators and such like may additionally be utilizedherewith. The central ball joint 100 is attached directly to aircraftstructure such as the pylon 12. The first sliding ball joint 102A isattached to the first A-arm 88A and the second sliding ball joint 102Bis mounted to the first A-arm 88A. The first and second sliding balljoint 102A, 102B permit sliding movement of the first and second A-arm88A, 88B (illustrated by arrow S in FIGS. 5A and 5B) to assure that onlya vertical load is reacted by the wiffle tree assembly 92. That is, thewiffle tree assembly 92 allows all engine thrust loads to be equalizedtransmitted to the engine pylon 12 through the rear mount platform 90 bythe sliding movement and equalize the thrust load that results from thedual thrust link configuration. The wiffle link 98 operates as anequalizing link for vertical loads due to the first sliding ball joint102A and the second sliding ball joint 102B. As the wiffle link 98rotates about the central ball joint 100 thrust forces are equalized inthe axial direction. The wiffle tree assembly 92 experiences loadingonly due to vertical loads, and is thus less susceptible to failure thanconventional thrust-loaded designs.

The drag link 94 includes a ball joint 104A mounted to the thrust case52C and ball joint 104B mounted to the rear mount platform 90 (FIGS.4B-4C). The drag link 94 operates to react torque.

The aft mount 84 transmits engine loads directly to the thrust case 52Cand the MTF 70. Thrust, vertical, side, and torque loads are transmitteddirectly from the MTF 70 which reduces the number of structural membersas compared to current in-practice designs.

The mount system 80 is compact, and occupies space within the corenacelle volume as compared to turbine exhaust case-mountedconfigurations, which occupy space outside of the core nacelle which mayrequire additional or relatively larger aerodynamic fairings andincrease aerodynamic drag and fuel consumption. The mount system 80eliminates the heretofore required thrust links from the IMC, whichfrees up valuable space adjacent the IMC 48 and the high pressurecompressor case 50 within the core nacelle C.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The disclosedembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

What is claimed is:
 1. A gas turbine engine comprising: a fan section; afirst compressor section including between four (4) and eight (8)stages; a second compressor section including between eight (8) andfifteen (15) stages; a first turbine section operable for driving thefirst compressor section, the first turbine section including betweenthree (3) and six (6) stages; a second turbine section operable fordriving the second compressor section; and a gear train defined along anengine centerline axis, wherein one of the first turbine section and thesecond turbine section is operable to drive the fan section through thegear train.
 2. The engine as recited in claim 1, wherein the firstturbine section includes three (3) stages.
 3. The engine as recited inclaim 2, wherein the first turbine defines a pressure ratio that isgreater than about 5.0.
 4. The engine as recited in claim 1, including afan case circumscribing the fan section that defines a fan bypassairflow, and wherein a bypass ratio of the engine is greater than about10.0.
 5. The engine as recited in claim 1, wherein said gear traindefines a gear reduction ratio of greater than or equal to about 2.3. 6.The gas turbine engine as set forth in claim 1, further comprising a fanvariable area nozzle to vary a fan nozzle exit area and adjust a fanpressure ratio of a fan bypass airflow of said fan section during engineoperation.
 7. The engine as recited in claim 6, further comprising: acontroller operable to control said fan variable area nozzle to vary afan nozzle exit area and adjust the fan pressure ratio of the fan bypassairflow.
 8. The engine as recited in claim 7, wherein said controller isoperable to reduce said fan nozzle exit area at a cruise flightcondition.
 9. The engine as recited in claim 7, wherein said controlleris operable to control said fan nozzle exit area to reduce a faninstability.
 10. The engine as recited in claim 6, wherein said fanvariable area nozzle defines a trailing edge of said fan nacelle. 11.The gas turbine engine as recited in claim 1, wherein the second turbinesection includes less than three (3) stages.
 12. The gas turbine engineas recited in claim 11, wherein a pressure ratio across the firstcompressor section and the second compressor section is greater thanabout fifty (50).
 13. The gas turbine engine as recited in claim 12,wherein the gear train comprises an epicyclic gear train with a gearreduction of greater than about 2.3:1.
 14. The gas turbine engine asrecited in claim 13, wherein the engine has a bypass ratio of greaterthan about 10.0.
 15. The gas turbine engine as recited in claim 14,wherein the fan section includes a plurality of fan blades and a fanpressure ratio across the plurality of fan blades is less than about1.45.
 16. The gas turbine engine as recited in claim 15, wherein acorrected fan tip speed of the plurality of fan blades is less thanabout 1150 ft/second.
 17. The gas turbine engine as recited in claim 15,wherein the fan blades are mounted to a fan rotor and the fan rotor issupported by a dual bearing.
 18. The engine as recited in claim 15,wherein the first turbine section defines a pressure ratio that isgreater than about 5.0.
 19. The engine as recited in claim 13, whereinsaid gear train defines a gear reduction ratio of greater than or equalto about 2.5.